Internal mixing of a portion of fan exhaust flow and full core exhaust flow in aircraft turbofan engines

ABSTRACT

A method of controlling plume exhaust heat and/or noise radiation from a turbofan engine assembly having a short nacelle. A mixer duct shell is supported such that a downstream edge of the short nacelle overlays an upstream portion of the mixer duct shell. A first portion of fan exhaust may be routed through the mixer duct shell between its inner surface and an outer surface of a core engine shroud. A second portion of fan exhaust may be routed over an outer surface of the mixer duct shell. At least one of the inner surface and an outer surface of the mixer duct shell may have an acoustic lining including a honeycomb core structure.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. application Ser. No.14/172,060, filed Feb. 4, 2014, which is a divisional of U.S.application Ser. No. 12/779,488, filed May 13, 2010, (now U.S. Pat. No.8,726,665), which is a continuation-in-part of U.S. application Ser. No.11/758,406, filed Jun. 5, 2007 (now U.S. Pat. No. 7,762,057). The entirecontents of the foregoing applications are incorporated by referenceinto the present application.

FIELD

The present disclosure relates generally to aircraft and, moreparticularly, to controlling noise and exhaust plume heat radiated fromaircraft turbofan engines.

BACKGROUND

The statements in this section merely provide background informationrelated to the present disclosure and may not constitute prior art.

In conventional aircraft, engines commonly are installed beneath theaircraft wings. Jet noise produced by the engines (most commonly, the“roar” heard at takeoff) can radiate largely unabated to nearbycommunities. Engine exhaust noise can be amplified when the sound isreflected off undersurfaces of the wing or other aircraft surfaces. Heatemitted from the engine exhaust plume also tends to be reflected off ofwings and pylons. In the case of military or derivative aircraft, thisreflected heat can tend to increase susceptibility of turbofan-poweredaircraft to heat-seeking missiles, when operated in threat environments.

SUMMARY

In one aspect the present disclosure may comprise a method ofcontrolling plume exhaust heat and/or noise radiation from a turbofanengine assembly having a short nacelle, and where the turbofan engineassembly is mounted on an airborne mobile platform. The method maycomprise providing a mixer duct shell supported such that a downstreamedge of the short nacelle overlays an upstream portion of the mixer ductshell. A first portion of fan exhaust may be routed through the mixerduct shell between an inner surface of the mixer duct shell and an outersurface of a core engine shroud. The core engine shroud may cover a coreengine. The first portion of the fan exhaust may be directed towards anozzle through which engine exhaust gasses pass. A second portion of fanexhaust may be routed over an outer surface of the mixer duct shell. Atleast one of the inner surface and an outer surface of the mixer ductshell may have an acoustic lining including a honeycomb core structure.

In another aspect the present disclosure may comprise a method ofcontrolling plume exhaust heat and/or noise radiation from a turbofanengine assembly having a short nacelle, and where the turbofan engineassembly is mounted on an airborne mobile platform. The method maycomprise providing a mixer duct shell supported such that a downstreamedge of the short nacelle overlays an upstream portion of the mixer ductshell. A first portion of fan exhaust may be routed through the mixerduct shell between an inner surface of the mixer duct shell and an outersurface of a core engine shroud, wherein the core engine shroud iscovering a core engine. The first portion of the fan exhaust may bedirected towards a nozzle through which engine exhaust gasses pass. Asecond portion of fan exhaust may be routed over an outer surface of themixer duct shell. At least one of the inner surface and an outer surfaceof the mixer duct shell may be covered with an acoustic lining. Theacoustic lining may include a plurality of distinct layers of material.

In still another aspect the present disclosure relates to a method forcontrolling plume exhaust heat and/or noise radiation from a turbofanengine assembly having a short nacelle, and where the turbofan engineassembly is mounted on an airborne mobile platform. The method maycomprise providing a mixer duct shell having an upstream portiondisposed adjacent to a downstream edge of the short nacelle. A firstportion of fan exhaust may be routed through the mixer duct shellbetween an inner surface of the mixer duct shell and an outer surface ofa core engine shroud. A second portion of fan exhaust may be routed overan outer surface of the mixer duct shell. At least one of the innersurface and an outer surface of the mixer duct shell may be covered withan acoustic lining. The acoustic lining may include a plurality ofdistinct layers of material that operate to absorb sound.

Further areas of applicability will become apparent from the descriptionprovided herein. It should be understood that the description andspecific examples are intended for purposes of illustration only and arenot intended to limit the scope of the present disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings described herein are for illustration purposes only and arenot intended to limit the scope of the present disclosure in any way.

FIG. 1 is a partial right rear perspective view of an aircraft andengine assembly in accordance with one implementation of the disclosure,with a portion of a mixer duct shell cut away;

FIG. 2A is a longitudinal sectional diagrammatic view of an engineassembly in accordance with one implementation of the disclosure;

FIG. 2B is a view of a leading edge of a mixer duct shell of the engineassembly shown in FIG. 2A;

FIG. 3A is a partial longitudinal sectional diagrammatic view of anengine assembly in accordance with one implementation of the disclosure,in which a thrust reverser is in a retracted position;

FIG. 3B is a partial longitudinal sectional diagrammatic view of anengine assembly in accordance with one implementation of the disclosure,in which a thrust reverser is in a deployed position;

FIG. 4A is a perspective view of an engine assembly in accordance withone implementation of the disclosure, in which forward and aft portionsof a mixer duct shell are contiguously positioned;

FIG. 4B is a perspective view of an engine assembly in accordance withone implementation of the disclosure, in which a movable aft portion ofa mixer duct shell is positioned apart from a fixed forward portion ofthe shell to provide maintenance access to underlying aft enginehardware while the engine is not in operation;

FIG. 4C is a perspective view of an engine assembly in accordance withone implementation of the disclosure, in which a movable aft portion ofa mixer duct shell is positioned apart from a fixed forward portion ofthe shell, and an assembly of one side of an aft part of a shortnacelle, core engine shroud and fixed portion of the mixer duct shellare raised together on a pylon mounted hinge and supported by attachedrods;

FIG. 5A is a partial side perspective view of an attachment device inengaged position in accordance with one implementation of thedisclosure;

FIG. 5B is a top view of the attachment device shown in FIG. 5A, theview taken along lines 58-58 of FIG. 5A;

FIG. 6 is a partial side perspective view of an attachment device indisengaged position in accordance with one implementation of thedisclosure;

FIG. 7 is a top view of the attachment device shown in FIG. 6;

FIG. 8 is a simplified side view of one specific configuration of aturbofan engine in accordance with the present disclosure;

FIG. 9 is a highly enlarged, cross sectional side view of a portion ofan acoustic lining that may be included on the inner surface, the outersurface, or both inner and outer surfaces of the mixer duct shell ofFIG. 8, and where the acoustic lining is shown formed by a pair offacesheets that may be arranged to sandwich a honeycomb coretherebetween;

FIG. 10 is a highly enlarged, cross sectional side view of another formof acoustic lining that may be used with the mixer duct shell of FIG. 8,where the acoustic lining is formed by a pair of facesheets that may bedisposed on one surface, or both opposing surfaces, of a honeycomb core,and with holes extending through the central honeycomb core of theacoustic lining;

FIG. 11 is a highly enlarged, cross sectional side view of another formof acoustic lining that may be secured to one or both opposing surfacesof the mixer duct shell, where an outer surface of the acoustic liningincludes a facesheet secured thereto and an inner surface forms a bulkabsorber layer, with a honeycomb core sandwiched therebetween;

FIG. 12 is a highly enlarged, cross sectional side view of another formof acoustic lining that may be secured to one or both opposing surfacesof the mixer duct shell, where the acoustic lining is formed by a pair15 of layers of bulk absorber material separated by a honeycomb centersection;

FIG. 12A is a simplified illustration showing how the acoustic liningmay preferably be tuned to attenuate specific types of noise present ineach of three distinct regions R1, R2 and R3 of a turbofan jet engine,with region R1 representing a region where fan noise is predominant,region R2 representing a region where low pressure turbine noise ispredominant, and region R3 representing a region where mixing noise ispredominant;

FIG. 12B, is a graph illustrating exemplary insertion loss in decibelsversus frequency for a double peak curve that represents a narrow lowfrequency band, where jet mixing noise is prevalent, and a broader highfrequency band where the turbine and fan noise collectively form asingle broader peak;

FIG. 12C is a simplified side view showing the upper and lowerbifurcations that may be used to support the mixer duct shell within theshort nacelle;

FIG. 12D is a simplified plan view looking down on the upper bifurcationshown in FIG. 12C, and illustrating the upper fairing that is used atthe leading edge of the mixer duct shell and upper bifurcationinterface;

FIG. 12E is a simplified front view looking at the fairing andbifurcation and mixer duct shell in accordance with directional arrow12E in FIG. 12D, and illustrating the fairings used at the intersectionof the upper bifurcation and the mixer duct shell;

FIG. 12F is a simplified side cross sectional side view showing a carrythrough element being used to extend through two acoustic liningelements placed on opposing surfaces of the mixer duct shell;

FIG. 12G is a simplified side cross sectional view of a carry throughelement extending through just one acoustic lining and being secured toan exterior surface of the mixer duct shell;

FIG. 12H is a simplified side cross sectional view of a carry throughelement extending through one acoustic lining and being secured to acountersunk fastening element in the mixer duct shell;

FIG. 12I is a plan view of a portion of the honeycomb structure of theacoustic lining showing how the corners are removed from intersectingcells to create a hole through which the carry through element mayextend;

FIG. 13 is a perspective view of just the mixer shown in FIG. 8;

FIG. 13A is a perspective view of a portion of the mixer shown in FIG.13 looking into the entrance of the mixer and into the plurality ofmixer flow channels;

FIG. 13B is a perspective view of a portion of the mixer of FIG. 13looking upstream from the aft end of the mixer;

FIG. 13C is a partial perspective view of a portion of mixer showing oneof the air flow vent paths formed therein for the engine caseventilation air, and another flow path formed therein for the engineshaft bearing cooling air;

FIG. 14 is a side view of the mixer of FIG. 13 that shows the mixerconfined by the contour of the mixer duct shell;

FIG. 15 is a view looking into the entrance end of the mixer of FIG. 13;

FIG. 16 is a view looking directly into at the exit end of the mixer ofFIG. 13;

FIG. 17 is a partial cross sectional view of just one of the integralvent paths taken in accordance with section line 17-17 in FIG. 16;

FIG. 18 is a side view of the mixer coupled to the aft end of the coreengine turbine frame and core exhaust case and also illustrating thevarious flows and how they are mixed and routed by the mixer;

FIG. 19 is a prior art simplified side cross sectional view of a thrustreverser in operation on a conventional turbofan jet engine and thetypical flows that exist during this phase of operation of the engine;and

FIG. 20 is a simplified side cross sectional view of the turbo fan jetengine of the present disclosure illustrating the various flows whenthrust reverser operation is occurring.

DETAILED DESCRIPTION

The following description is merely exemplary in nature and is notintended to limit the present disclosure, application, or uses.

An aircraft adapted in accordance with one implementation of thedisclosure is partially shown and indicated generally in FIG. 1 byreference number 20. The aircraft 20 has preferably one or more turbofanengine assemblies, one of which is shown in FIG. 1 and indicatedgenerally by reference number 24. The engine assembly 24 includes a coreengine 28 having an attached tail cone exhaust plug 32. The engineassembly 24 includes a short nacelle 36 mounted to a wing pylon 38. A“short” nacelle is one that is shorter than its associated core engine.Thus the short nacelle 36 provides a short fan duct 40 through whichexhaust from a fan forward of the core engine 28 may exit the nacelle 36alongside the shroud cover over the core engine 28.

One configuration of a mixer duct shell is indicated by reference number44. A portion of the shell 44 is shown as having been cut away topartially expose the core engine 28. The shell 44 is substantiallycoaxial with both the fan duct 40 and shroud cover of the core engine28. The shell 44 extends forwardly into the fan duct 40 to provide aninterstitial mixer duct 48 between the shell 44 and both the shroudcover of the core engine 28 and the tail cone exhaust plug 32. Asfurther described below, the fan duct 40 and interstitial mixer duct 48are configured to provide a means of mixing a partial (e.g., a minority)amount of fan exhaust with the core engine exhaust and to bypass theother (e.g., majority) amount of fan exhaust out the exit of nacelle 36alongside the outer surface of the mixer duct shell 44.

One configuration of an engine assembly is indicated generally in FIG.2A by reference number 100. A short nacelle 102 is mounted around aforward portion 104 of a core engine 106. The core engine 106 is coveredby a core engine shroud 108. The core engine 106 extends through and aftof the nacelle 102 and ends with a tail cone exhaust plug 110. Thenacelle 102 includes a fan case frame 190 and nacelle shroud 112 mountedon the core engine 106 by struts 116. Air enters the nacelle 102 througha turbofan inlet 118 and travels through a fan pressurized duct 120. Airalso enters core inlets 122 aft of a nose cone 192 and fan portion 124of the core engine 106. The nacelle 102 has an exit nozzle 126. In thepresent exemplary embodiment, an edge 128 of the nozzle 126 includes aplurality of chevrons 130. It should be noted generally that engineassemblies having other short nacelle configurations could be adapted inaccordance with principles of the disclosure. For example, in someimplementations, a nacelle exit nozzle may include a variable areanozzle and not include chevrons.

One configuration of a mixer duct shell 134 is mounted on an aft portion140 of the core engine 106. A forward portion 144 of the shell 134extends forwardly into the nacelle exit nozzle 126 to provide aninterstitial mixer duct 148 between the shell 134 and the shroud 108extending over the core engine aft portion 140. A leading edge 182 ofthe shell 134 is shown in greater detail in FIG. 28. The radius of theleading edge 182 is tailored for aerodynamic contouring that preventsflow separation that would otherwise cause a reduction in fuelefficiency and increased noise. The shell forward portion 144 is affixedto the core engine shroud 108 by a plurality of circumferentially spacedand aerodynamically tailored radial pillars 150. The pillars arepreferably oriented in line with the crowns of a lobed mixer 164 tominimize noise. An aft portion 152 of the shell 134 may be moved in anaftward direction along a pillar slide 154 with weight supported on asliding track 184 attached to an engine pylon sidewall 186. Moving theaft portion 152 provides access to underlying structure, e.g., access toportals for turbine inspections and access to sections of the shell 134and aft core engine 140. An exemplary range of possible translation ofthe aft portion 152 is indicated by an arrow 156. The movable shellportion 134 has a convergent exit nozzle 158. In the present exemplaryconfiguration, the nozzle 158 includes a plurality of chevrons 160 alonga nozzle edge 162. Other or additional fixed or variable area nozzleconfigurations are contemplated, however, that do not include chevrons.

The lobed mixer 164 is positioned aft of a turbine frame section 166 ofthe core engine 106 and may be integrated with the tail cone exhaustplug 110. The lobed mixer 164 is high penetration to enhance mixing.Less penetrating mixers can also be incorporated as to optimize mixingversus duct-path flow losses. While lobed mixer prior art shows lessthan about 65% penetration due to greatly increased costs and diminishedperformance/noise payoffs, the inner duct height of the exemplary art isrelatively small in comparison, so achieving almost 100% penetration isboth desirable and not relatively cost prohibitive. The lobed mixer 164may be scalloped as dictated by an optimum contour for the engine. Themixer 164 is positioned inside the shell 134 and upstream of the shellexit nozzle 158. It should be noted that configurations are contemplatedin which other or additional means of mixing may be provided upstream ofthe mixer duct shell nozzle exit 158. Configurations also arecontemplated in which no lobe mixer is provided.

In various configurations, the shell 134 is sufficiently cooled by thefan airflow during engine operation such that lower cost liners foracoustic absorption and/or debris containment may be structurallyintegrated into the engine assembly at minimal weight. Debriscontainment liners may be made, e.g., of woven composite. In variousengine assembly configurations, various liners may be provided forreducing noise and/or for containing turbine fragments. For example, theengine assembly 100 includes an acoustic lining 168 covering as much ofan inner surface wetted area 170 as practical. Acoustic lining 168 alsomay be provided, e.g., along part of an outer surface 172 of the shellfixed forward portion 144 opposite the nacelle inner surface 170, and onportions of a surface 146 of the core engine shroud 108. Acoustic lining168 also may be provided, e.g., along part of an aft inner surface 174of the shell movable aft portion 152 and on a conical trailing surface114 of the exhaust plug 110. In some applications, the acoustic linerson the aft inner surface 174 and conical exhaust plug surface 114 aretuned for control of jet mixing noise. Containment liners 176 may beprovided in the vicinity of the turbine frame section 166, e.g., alongan inner surface 178 of the shell fixed forward portion 144 and along aforward inner surface 180 of the shell movable aft portion 152. Itshould be noted that acoustic and/or containment liners may be providedin various ways and locations, or not provided at all, dependent, e.g.,on engine assembly structure and performance.

In some embodiments, the leading edge 182 of the shell 134 penetratesinto the nacelle exit nozzle 126 sufficiently far to act as anaerodynamic inlet to the mixing duct 148 at fan stage pressure. The coreengine shroud 108 may be re-contoured so as to allow penetration of theleading edge 182 to be for a minimal longitudinal length forward fromthe fan nozzle exit edge 128. For retrofit applications the core engineshroud 108 surface may be re-contoured inward opposing the thickenedleading edge 182 to accommodate increased leading edge thickness withoutreducing cross sectional area for the portion of fan flow entering theinterstitial mixing duct 148. Keeping a constant inlet cross-sectionalarea ensures that the entrance velocity is minimized and velocitydependent pressure losses are lowered as the fan air moves down theinterstitial duct toward the mixer. In new engine applications, normaldesign methods will ensure that the duct work accommodates the correctflow path areas, at entrance, along the paths and at exit to optimizeperformance over the range of powered conditions. In generally allapplications the mixer duct leading edge 182 is also positioned aft of amechanism for thrust reversal and does not affect or inhibit thrustreverser function, reliability or reverse thrust generation. In someapplications the core barrel shroud 108 has acoustic lining covering asmuch of the flow-wetted area as practical.

An exemplary engine assembly is partially shown and indicated generallyin FIGS. 3A and 38 by reference number 200. The assembly 200 includes ashort nacelle 202. Referring to FIG. 3A, a thrust reverser 204 is in aretracted position. Fan exhaust 208 flows along both sides 212 and 216of a mixer duct shell 220. Engine core exhaust 224 flows from a mixer228 and is mixed internally with fan exhaust 208. Mixed exhaust 232exits the engine assembly through a nozzle 236 of the shell 220.Referring to FIG. 38, as a rear portion 240 of the nacelle 202 slidesaft and, e.g., exposes “C shaped” cascade type reverser vanes 244, thereare no interferences with the shell 220. During normal reverseroperation, core flow 224 exits the mixer 228, but without interleavedfan flow, and thus tends to slow down (diffuse) through the nozzle exit236—at a lower velocity than typically would be the case in the absenceof the shell 220. A gap 250 between a deployed diverter gate 254 of thethrust reverser 204 and leading edge 258 of the mixer duct shell 220determines a residual amount of cool air flow entrained backwards fromthe fan exhaust nozzle exit into the interstitial mixer duct by ejectioneffect of the diffused core exhaust flowing through the lobed mixer.This gap can be optimized, for a given implementation, to createadequate diffusion of core exhaust and sufficient backflow entrainmentfrom the fan exit to keep mixer surfaces cooled and to lessen theresidual forward thrust component. Thus, advantageously, net magnitudeof reverse thrust (thrust reverser effectiveness) can be increased, forimplementations with at least some types of thrust reversers.

In various implementations of the disclosure, an apportioned amount ofpressurized fan air combined with core exhaust flow, e.g., a 2:1 massflow ratio, can be captured and ducted. In typical current generationhigh-bypass engines (past generation were substantially less than 2:1)where the total fan air mass flow to turbine core flow ratio ranges fromabout 4.5:1 to greater than about 10:1, the apportioned mass flow ratioof 2:1 represents a minority portion of the fan duct mass flow. At abouta 2:1 mass flow mixing ratio, a prior art scalloped lobe mixer can beused, e.g., to internally mix core exhaust to a core static temperatureat least 200 degrees K less than the temperature in an engine assemblyin which such mixing is not performed and when the engine is at nearlyfull take-off thrust or climb thrust settings. Additionally in highthrust conditions, the 2:1 mixed core exhaust may exhibit a corevelocity at least 250 feet per second less than the core velocity of anengine assembly in which mixing is not performed.

As previously mentioned, for maintenance access, a mixer duct shellmovable portion can be detached from an affixed portion of the shell andslid aft. One configuration of an engine assembly is indicated generallyin FIGS. 4A, 4B, and 4C by reference number 300. The assembly 300includes a short nacelle 304 and a mixer duct shell 308 having a fixedforward portion 312 and a movable aft portion 316. As shown in FIG. 4A,the forward and aft portions 312 and 316 are contiguous. As shown inFIG. 4B, the aft portion 316 is extended away from the forward portion312 to expose part of an engine core shroud 320. As shown in FIG. 4C,for some turbofan engines, access can be gained to the engine corecasing and components underneath an aft section 306 of the nacelle 304and underneath the engine core shroud 320, forward of the lobed mixer322. For such types of engines, the engine core shroud 320, the fixedportion of the mixer duct shell 312, and the aft section 306 of the fannacelle 304 are joined as one duct assembly 330 on each side of theengine, forming two halves each hinged at the upper pylon seam and splitat the lower centerline. The halves are detachable along a lower seamsupported by structural frame member and lower bifurcation 326 extendingaft from the forward fan case cowl and fan case frame 328. The ductassembly 330 can be raised and supported by rods 332 employing standardprocedures to gain full access to the core engine, after the movableportion of the mixer duct 316 has been suitably detached and translatedback, out of the way.

A forward portion of a mixer duct shell is fixedly supported from theforward parts of circumferentially spaced radial pillars that extendfrom the core engine shroud outward to an interior concentric surface ofthe mixer duct shell. One such pillar attachment device is indicatedgenerally in FIGS. 5A-7 by reference number 400. A base surface 416 of aforward radial pillar part 410 is fixedly attached to the core engineshroud and the forward mixer duct shell is fixedly attached to anoutward surface 414 of the forward radial pillar part 410. A movablemixer duct shell aft portion is fixedly attached to an outward surface424 of an aft movable pillar part 420 of the attachment device 400. Aplurality of circumferentially-spaced forward pillar parts 410 areattached to the mixer duct shell's forward portion to provide for itsradial standoff from the engine core shroud, forming an interstitialmixer duct 148, shown in FIG. 2A. An equal plurality of aft pillar parts420 are attached in clocked alignment with forward pillar parts 410. Theaft pillar parts 420 slide longitudinally along dovetail keyed surfaces432 built into the forward pillar parts 410. The keyed surfaces act aslongitudinal guides and serve to further constrain azimuth clockingmotions that could otherwise compromise the integrity of the shell partlatch-up. The weight of the aft mixer shell and attached pillar aftparts 420 are supported by a rail slide 184 built into the sidewall ofthe engine pylon 186, as previously referenced in FIG. 2A. As shown inFIG. 6, the pillar parts 420 separate completely from their dovetailkeyed sliding surfaces 432 as the movable portion of the mixer ductshell is slid aft ward. When slid forward back together as shown in FIG.5A, the keyed sliding surfaces re-engage, thereby holding the slidingparts into alignment. Thus the attachment devices 400 permit the mixerduct shell aft portion to be released and slid longitudinally aft-wardquickly with full weight supported by the sliding track on the pylon andthen slid back together quickly and re-engaged securely with a latchmechanism 430 built into pillar forward part 410.

When the movable portion of the mixer duct shell is slid intoengagement, that is, when the pillar portions 420 and 410 are positionedinto engagement, a rotating lock rod 458 with a hex (or alternate) headmay be turned ninety degrees with a hex socket tool to securely lockeach movable radial pillar part 420 to its adjacent fixed portion 410.FIG. 5B indicates the motion of a hooked latch 450, doweled catch 428,cam 412 on a lock rod 458 and a spring loaded pin 404 that addsresistance to the latch 450 while it is in the locked position so as tokeep it held in tension on the catch 428. When the lock rod is rotatedback for release as shown in FIG. 7, the spring loaded pin 404 and newposition of the cam 412 cause the latch to spring open and remain open,freeing the movable portion 420 to slide aft. Thus the movable portion420 may be secured or released by means of a plurality of (for example,eight) ninety-degree twists of the rotating lock rods 458. Supportingradial pillar portions 410 and 420 may be circumferentially spaced incoincidence with crowns of the lobes of a flow mixer, for example, themixer 164 shown in FIG. 2A. Alignment of pillars 400 with mixer lobesreduces surface Mach numbers on the downstream lobes and reduces overallviscous losses.

Various implementations make it possible to incorporate additionalacoustic absorbing liners and/or chevron nozzle treatments if desired.Acoustic liners may be, for example, porous honeycomb sandwich-styleacoustic absorbers or bulk absorbing material manufactured frommetallic, ceramic or composite material in a foam-like or generallyporous fabrication. Such absorbers may be used, for example, to line atleast a portion of a forward fan cowl and aft fan nozzle. Someimplementations may include opposing acoustic liners and a linerstructurally disposed on an outer surface of an affixed portion of thecore engine shroud, mixer duct shell, inner surface of the mixer duct,and tail cone exhaust plug such as previously described with referenceto FIG. 2A. Such linings may cooperate with a fan exit nozzle liner toabsorb modal and broadband noise propagating aft-ward through enginebypass ductwork.

Referring again to FIG. 2A, the inner surface of the shell movableportion 152 near the aft mixed-jet nozzle 158 may also be structurallydisposed with additional acoustic absorber lining. Such lining candiminish sound intensity of modal and/or broadband noise generated bythe fuel combustor, rotating turbo-machinery, stator vanes, turbulentflow mixing and mixer trailing edges. In some configurations,application-specific, supplemental low-to-medium temperature-tolerantwoven polymeric liners may be attached to inward facing surfaces of themixer duct shell affixed portion 144 and forward segment of the innersurface of the movable portion 152, upstream of the lobed mixer 164. Insuch manner, containment may be enhanced, in an azimuthally directionalfashion, of any hot turbine machinery debris that might be released andthat might be capable of penetrating through the engine encasement,e.g., during destruction of the aft turbine area by guided missileimpact. Directional characteristics of containment may be attained byplacement of a containment liner with such a sub-tense of arc, thatfragments that otherwise would impact and possibly penetrate wing skin,wing structure, control surfaces, engine support structure, or fuselagewould have sufficient energy dissipated by the added protection layer.Advantageously, supplemental turbine-stage containment liners would inno way alter performance of normal fan-stage containment features.

Referring again to FIG. 2A, the outer core engine barrel 108 may bere-contoured to permit reduction of flow path losses from theinterstitial mixer duct inlet through the surfaces of the lobed mixerand in application-specific configurations have additional acousticlining, preferably encompassing as much of the wetted area as practical.The duct shell leading edge 182 is aerodynamically contoured to preventseparation and may be shaped to act as a low-loss flow inlet and may bepositioned forward of the convergent section of the fan nozzle 126. Insuch manner, incoming gas velocity begins at a low fan duct velocity andpressure losses to the downstream mixer plane are minimized. Such achanneled gas path into the mixer duct 148 may remain diffused viasubtle increase of the cross sectional area from the inlet area toslightly larger area as permitted by re-contouring the surface of theopposing core engine shroud 108. This may serve to further accommodatethe radial thickness and outer flow line contour around the mixer ductshell 134 with minimal pressure loss in the contained flow path. Theincreased cross section may be retained through the fluid mixing devicebounded by the inside of the mixer duct shell 134. Flow path tailoringis accomplished via the above re-contouring of the core engine shroud108, shaping the mixer 164, and replacing the standard core nozzle withone of appropriately larger diameter. The contained flow path may remainadequately diffused at lower velocity until mixing is accomplished andthe mixed outflow enters the convergent end of the nozzle exit 158.

The mixer 164 may be tailored for a low expansion rate and small coreflow diffusion to accommodate lower scrubbing losses through the lobes.Lobe leading edge shapes may be broad, smooth and gradually tape, red toprecipitate a minimum in flow friction loss. Weight of the mixer andareas of surface affecting the scrubbing loss through the lobes may bereduced through scalloping the mixer's radial side walls. Scalloping ofthe mixer sidewalls enhances mixing prior to entering the convergentnozzle.

The above-described mixing can lower velocity differences and turbulentshear between mixer duct and fan bypass flows, resulting in much lowernear-field and far-field radiated jet noise. The mixing function alsocan significantly reduce aircraft plume radiance, thereby loweringsusceptibility and increasing ability to defend with onboard defensivejamming systems reliant upon IR sensors that view outward to pointtoward and track the threat—potentially through own-plume selfobscurations. In cases where there may be additional desire to reducevisibility or apparent temperature of hot metal exhaust. parts, suchreductions may be accomplished within the scope of this disclosure usingvarious implementations of alternative mixer device designs: ones thatimpinge more cool airflow onto the aft exhaust plug surface, ones thatutilize commercial or non-commercial low-emissivity coatings, and onesthat twist the mixer surfaces so as to inhibit visibility to hightemperature turbine areas and guide vanes. Various implementations ofthis disclosure are contemplated to employ any of a variety of mixerdesigns, or no mixer at all, to accomplish the reduction of radiatedheat.

The structural integration of acoustic liners and debris containmentliners can enhance reduction of interior engine cavity noise and canprovide additional containment barriers for impact-generated debris,e.g., in cases where the aircraft engine cannot be defended againstimpact by IR guided missiles. The mixer duct shell attachment structureserves to permit support of the weight of the movable portion of theshell and facilitates its movement aft-ward to expose interior surfacesand parts for easier maintenance. In addition, the attachment structurefacilitates moving the mixer duct shell aft-ward, thereby allowing useof normal maintenance procedures on modern turbofan engines that raiseand support nacelle assembly halves to expose the core engine for easymaintenance. The affixed portion of the mixer duct inlet is positionedsuch that no interference is generated with the mechanisms and functionsof the thrust reversers, and is also uniquely tailored to not induceself-noise or entrance flow losses. The above flow path features,including diffusion and flow shaping, can reduce flow path losses in theinterstitial mixer duct up to and through the radial blending of the twoflows, prior to entrance into the convergent mixed nozzle exit.

The foregoing discussion can be seen to describe a method forcontrolling exhaust plume heat and noise radiation from an aircraftturbofan engine assembly. A first portion of fan exhaust is routedthrough an interstitial mixer duct formed between an inner surface of amixer duct shell and a core engine shroud of the engine assembly to anozzle through which engine exhaust passes. A second portion of fanexhaust is routed over an outer surface of the mixer duct shell.

A nacelle mixing design implemented in accordance with the disclosurepromotes internal mixing and can substantially slow jet velocities atthe exit of an aircraft engine. Various implementations cansimultaneously reduce take-off and landing airport community aircraftnoise, cruise aft-cabin shock-cell noise and infrared plume emissions ofturbofan powered aircraft. The foregoing engine assembly can locallyabsorb additional damage that might be induced by high-velocity debrisreleased into areas surrounding the aft engine system should the aftengine portion on an aircraft be impacted, e.g., by a heat seekingmissile, protecting against possible collateral damage to fuel tanks,wing structures, control surfaces and/or fuselage.

Referring to FIG. 8, one specific configuration of another turbofanengine assembly 100′ in accordance with the present disclosure is shown.The various components shown in FIG. 8 that are common with thosediscussed in connection with FIG. 2A have been denoted with the samereference numbers as used in the discussion of the embodiment of FIG.2A, but with the reference numbers used in FIG. 8 also including theprime designation (i.e., “′”). It will be appreciated that thecomponents described in connection with FIG. 2A apply for the turbofanengine assembly 100′ with the exception of a new construction for themixer duct shell 134′ and the incorporation of a new center bodyventilation and mixer device 194, which will hereafter be referred tosimply as the “mixer 194”. The mixer duct shell 134′ can be seen to besupported by a pillar slide 184′ such that a leading edge 182′ of themixer duct shell 134′ extends within an exit nozzle 126′ of the shortnacelle 102′. A downstream edge 162′ of the mixer duct shell may includechevrons 160′.

With further reference to FIG. 8, instead of separate components thatform the tail cone exhaust plug 110 and the lobed mixer 164, theturbofan engine assembly 100′ incorporates the new mixer 194. The mixer194 is disposed at a downstream side of a core engine 106′ having a fan124′ and mounts to the aft portion 140′ of a turbine case 141′ and theaft portion of a core exhaust shroud (as also noted in FIG. 14). Themixer 194 is substantially entirely surrounded by the mixer duct shell134′. The mixer 194 and its function will be described in greater detailin the following paragraphs and in connection with FIGS. 13-20.

Referring now to FIG. 9, a highly enlarged cross sectional portion of anacoustic lining 137 that may be secured to an inside surface HOP Ref.7784-001052/DVC and/or an outside surface of the mixer duct shell 134′in shown. In this embodiment the acoustic lining 137 may incorporate acontinuous, high temperature, composite honeycomb core structure,indicated by number 137 a. General material compositions for thecomposite core can include light-weight ceramics, ceramic matrixcomposites (CMCs) or porous composites to save weight.

A first facesheet 137 b may be secured, either by weld adhesives orother suitable means, to an exterior surface 137 c of the honeycomb core137 a. Similarly, a second facesheet 137 d may be secured by weldadhesives or other suitable means to the inner surface 137 e of thehoneycomb core 137 a. Each of the first and second facesheets 137 b and137 d may have a plurality of perforations or pores 137 f and 137 gformed therein. The facesheets 137 b and 137 d may each be made fromhigh temperature materials including titanium and titanium alloys,Inconel, or CMC for the inner (high temperature) facesheet, and similarmaterial choices or lighter metal alloy liners used in commercial enginefan ducts for the outer (lower temperature) facesheets. Temperaturetolerance and cost dictate the optimized solution for each acousticliner application. The facesheets 137 b and 137 d for the acousticlining 137 may have a preferred thickness of substantially 0.004 inch(0.1 mm) in regions of active acoustic absorption. The thickness mayvary in the presence of joints, attachment points, or other regionsnecessitating structural strength. The latter could for example be in aregion of penetrations for systems housings, vents, or otherattachments. However, it will be appreciated that this thickness mayvary somewhat depending on the specific material that is used for thefacesheets 137 b, 137 d and the precise location on the mixer duct shell134′ that the facesheets 137 b, 137 d are located at. The shape of theperforations may be circular or otherwise substantially symmetric inshape, and the diameter or characteristic dimension preference variesdepending on whether the facesheet 137 b or 137 d covers the honeycombcore 137 a or forms a bulk absorber facesheet layer. Bulk absorberfacesheets 137 i 1 and 137 j 1 are shown in FIGS. 11 and 12. A preferreddiameter for the perforations 137 f, 137 g, when the facesheets 137 b,137 d are applied over the honeycomb core 137 a, may vary substantiallybetween about 0.016 inch and about 0.028 inch (0.4 mm-0.7 mm). Apreferred diameter for the perforations 137 i 1, 137 j 1 for the bulkabsorber facesheets 137 i, 137 j may vary substantially between about0.02 inch to about 0.16 inch (0.5 mm-4 mm).

The percent open area, or porosity (both are acceptable terms of art)that the perforations 137 f, 137 g form on the facesheets 137 b, 137 dmay vary substantially between about 5% and about 30%, and for a bulkabsorber facesheet substantially between about 25% and about 50%. Thedepths of the perforations 134 f and 134 g are preferably tuned for thediffering gas temperatures (i.e., speed of sound differences) on bothsides of the mixer duct shell 134′ and the differing spectral noisesignatures on both sides of the mixer duct shell 134′. The depths of theperforations 137 f and 137 g may be the same or may differ. Thethickness of the honeycomb core 137 b may be on the order of about 0.4inch-2.0 inches (10 mm-50 mm).

Typically, the depth of the honeycomb core 137 a will vary slightly overthe axial length of the mixer duct shell 137 a, with the depth of thehoneycomb core 137 a at the downstream edge of the mixer duct shell 162′tapering down to a depth which is less than that at the remainder of themixer duct shell 134′. It is an advantage that the first facesheet 137 band the second facesheet 137 d may include different thicknesses,different sized perforations 137 f and 137 g, as well as differentdensities of the perforations 137 f and 137 g. The size and densities ofthe perforations 137 f and 137 g may thus be tailored to meet the needs(i.e., differing gas temperatures and spectral noise signatures) onopposing sides of the mixer duct shell 134′ when the mixer duct shell134′ is used with a specific turbofan engine. Each turbofan engine willhave slightly unique characteristics for which each of the parameterscan be knowledgeably manipulated by someone skilled in this art of jetengine noise reduction.

Referring to FIG. 10, another embodiment of the mixer acoustic lining137′ is shown that includes the facesheets 137 b and 137 d, and inaddition includes holes 137 h formed through the honeycomb core 137 a.The holes 137 h extend completely through the honeycomb core 137 a. Theholes 137 h may vary in diameter, but may be typically within a diameterrange from the same as the smaller of the two opposing face sheet 137 b,137 d perforation diameters but not larger than the larger of the twoopposing face sheet perforation diameters in diameter, and range indensity within a range similar to that given above for the percentageopen area for the facesheets 137 b, 137 d. The holes 137 h act tostimulate comparable sound attenuation to a double-layer acousticlining, which may produce even more bandwidth covering. The embodimentof FIG. 10 is expected to be particularly useful for some engine cyclesand center-body designs.

Referring now to FIG. 11, another embodiment of the mixer duct shell137″ is shown that incorporates a bulk absorber facesheet 137 i on theinner surface 137 e (i.e., the hot gas side of the mixer duct shell134′) rather than the facesheet 137 d. The bulk absorber facesheet 137 iis especially advantageous for applications where the structural depthof the downstream edge of the mixer duct shell 134′ is too thin toeasily incorporate the second facesheet 137 d. The bulk absorberfacesheet 137 i may vary in thickness but is typically within about 0.4inch-1.0 inch (1 Omm-25 mm) thick, and is secured to the inner surface137 e of the honeycomb core 137 a by weld adhesives or other suitablemeans. The bulk absorber facesheet 137 i may include perforations 137 i1 having dimensions and densities described above for the perforations137 f and 137 g. The bulk absorber material 137 i may be preferred foruse in applications where the structural depth of the honeycomb core 137a is limited and the hot stream (i.e., the stream on the inner surface137 e side) has more mixing noise than other sources in the acousticsignature.

Referring to FIG. 12, still another embodiment of the acoustic lining137′″ is shown. In this embodiment the acoustic lining 137′″ includestwo bulk absorber facesheets 137 i and 137 j, with the bulk absorber 30facesheet 137 i including perforations 137 h and bulk absorber facesheet137 j including perforations 137 j 1. Bulk absorber facesheet 137 i issecured to the inner surface 137 e of the honeycomb core 137 a whilefacesheet 137 j is secured to the outer surface 137 c. Each of the bulkabsorber facesheets 137 i and 137 j may have similar or differingthicknesses, but in most instances will be similar to the thicknessesgiven above for facesheet 37 i. Alternatively, one or both of the layers137 i and 137 j may be provided without perforations. The use of bulkabsorber facesheets 137 i and 137 j on both sides of the honeycomb core137 a may be preferred when the structural depth of the honeycomb coreis limited on both sides.

With reference to FIG. 12A, it will also be appreciated that the mixerduct shell 134′ may incorporate two or more different ones of theacoustic lining 137 constructions explained in connection with FIGS.9-12 at different areas of the mixer duct shell 134′. It will beappreciated that for most High Bypass Ratio (HBR) turbofan engines, theacoustic field encountered by acoustic linings in the upstream region ofthe mixer duct shell 134′ (region R 1) is aft-radiated fan noise. Aunique factor with the inner surface of the mixer duct shell 134′ isthat it will also be exposed to low pressure turbine noise which willexist in region R2. Accordingly, for at least the forward one-third ofthe mixer duct shell 134′, the properties of the acoustic lining on boththe outer and inner side of the mixer duct shell may be substantiallythe same, with exception of differences (if significant) between thelocal speed of sound. Due to heat soaking, the inner part of the forwardmixer duct shell 134′ likely encounters a slightly higher temperatureand corresponding speed of sound. For the roughly the middle third ofthe inner mixer duct shell (region R2), the dominant tonal sound fieldis generally from the low pressure turbine. The last third (region R3)is dominated by jet mixing noise. Thus, the acoustic lining on the innersurface of the mixer duct shell 134′ may be divided and tailored intosubstantially three acoustic treatment zones: 1) fan (region R1); 2)turbine (region R2) and 3) jet mixing noise (region R3). Conversely, theprimary tuning priority of the acoustic lining on the outer surface ofthe mixer duct shell is fan tonal sources. It is also preferred that theacoustic lining in each region be manufactured as a continuous 360degree barrel section.

The primary role of the forward/upstream part of the acoustic lining 137on the outer wall of the mixer duct shell 134′ will be targeting thesame sound source(s) that the acoustic lining on the fan duct inner wall36 a′, as shown in FIG. 18. Because the annular fan “duct” terminates atcompletion of the fan-nozzle trailing edge (or Chevrons, if they areused) the acoustic lining 137 on the outer wall of the mixer duct shell134′ can either be unchanged relative to the rest to save costs.Alternatively, the acoustic lining 137 downstream of the fan nozzle exiton the outer surface of the mixer duct shell 134′ can be altered totarget broadband noise through use of the same bulk absorbing materialused on the inner wall of the mixer duct shell 134′, and the same facesheet used on the forward part of the outer shell, but with a relativelyhigh level of porosity (25% to 40%).

With regards to acoustic lining on the inner surface of the mixer ductshell 134′ the following motivations for different special liningarrangements are articulated. Because airflow produced by the fan 124′(FIGS. 8 and 18) is split into first and second flows; source noiseenergy comparable to that which propagates on the outer side of themixer duct shell 134′ propagates on the inner side as well. The modalcontent will differ due to the presence of the mixer duct shell 134′.However, the multi-segment acoustic lining design presented in thepresent disclosure is expected to compensate for all of the significantmodes that contribute to far-field noise. The forward (upstream) mostregion of the acoustic lining 137 (i.e., Region 1) on the inner surfaceof the mixer duct shell 134′ may thus be preferably designed for thesame target tonal and broad-band frequencies as the inner surface of theouter fan duct for both cases where the mixer duct shell is originalequipment or a retrofit. This “fan source” zone extends backward tosubstantially the plane coincident with the fan nozzle trailing edge.The Low Pressure Turbine (LPT) zone (Region 2) is characterized byslightly higher tonal frequency noise with a significant broad-bandenergy covering a large frequency range. For Region 3, the preferredcavity depth will be less and the perforate diameter smaller(proportionate to dominant sound frequency and 30 the speed of sound byHelmholtz law—which is well known to those of skilled in the art), andthe ranges for these parameters have been previously explained herein))for the same face sheet thickness. This zone extends aft to the mixingplane. For the jet mixing zone (Region 3), most people skilled in theart are likely to presume that the desired jet noise frequency toattenuate is the typical “mixer lift” frequency observed in thefar-field which can occur between 4,000 and 5,000 Hz. Tuning theacoustic liner 137 for 4 KHz to 5 KHz, however, will not likely resultin effective far-field noise reduction however because this lift istypically the result of scattering. Accordingly, for the greatestoverall reduction in far-field noise an optimum target frequency for themixing noise is substantially about 2,000 Hz. To absorb this jet mixingsource as well as other modes from upstream sources that are notentirely absorbed by the upstream acoustic liner sections 137, thehoneycomb 137 a cavities may be replaced with the CMC bulk liningmaterial. The bulk lining material produces effective sound absorptionabout a highly broad range of frequencies about a central preferredtarget frequency of desire (i.e., about 2000 Hz).

It has been depicted that at least three unique zones {i.e., Regions1-3) are preferably employed for the acoustic lining 137 on the innersurface of the mixer duct shell 134′. It is furthermore apparent thatthese zones are preferably segregated in some way. For efficientmanufacturing it is preferred that the acoustic honeycomb linings 137 ofthe mixer duct shell be manufactured as complete elements as much aspossible. The result is therefore typically cylindrical or conicalhoneycomb acoustic ling elements of substantially uniform properties.Once assembled/fastened onto the mixer duct shell 134′ structure, it isthen preferred that both inner and outer facesheets, for examplefacesheets 137 b and 137 d, each be of a single piece construction, withthickness and perforations being specific to each zone. Any type ofinternal barrier between zones is not preferred except between honeycomband bulk material, with exceptions being possible for containment offluid isolation.

FIG. 128 is a graph of an exemplary in-duct “insertion loss” that may beprovided by the various embodiments of the acoustic lining discussedabove in connection with FIGS. 9-12. It will be appreciated that theterm “insertion loss” is a term of art, and that for the specificacoustic lining configuration that is implemented, the frequencyspecific Insertion illustration in FIG. 12B may be preferred.Specifically, a desired insertion loss will comprise a unique narrow andbroad double hump performance illustrated by humps 137K and 137L of thewaveform shown in FIG. 12B. A moderate frequency hump (i.e., hump 137K)is tailored toward jet mixing noise. A broad high frequency hump (hump137L) is actually the product of two adjacent discrete humps and istailored toward the fan and low pressure turbine (LPT) zonesrespectively. A tolerance band 137 m is shown in shading around anominal curve 137 n and represents acceptable unique optimizationopportunities for different engine designs.

With reference to FIG. 12C, it will be seen that upper and lower,radially inwardly extending bifurcations 81 and 82 are used to supportthe short nacelle 36′. FIG. 120 illustrates the upper bifurcation 81,and it can be seen that the upper bifurcation makes use of asemi-bulbuous, aerodynamically shaped fairing F1 where the upperbifurcation 81 meets the mixer duct shell 134′. FIG. 12E illustratesthat another semi-bulbuous fairing F2, identical to fairing F1, is alsoused on the opposite surface of the mixer duct shell 134′. It will beappreciated that the lower bifurcation 82 similarly includes a pair offairings F1 and F2 where it intersects the mixer ductshell 134′.

Referring now to FIGS. 12F-12I, it can be seen how a carry throughcomponent 135 may be used to extend through one or more of the acousticlining 137. The carry through component 135 may be a solid or possiblyeven tubular structural component, preferably made from metal, that maybe used for securing the acoustic lining(s) 137 on one or both sides ofthe mixer duct shell 134′. Referring specifically to FIG. 12F, the carrythrough component 135 may include an elongated element 135 a with areceptacle 135 b and 135 c at its opposite ends. The receptacles 135 band 135 c form attachment elements that may be welded or bonded to thefacesheets 134 b and 13 d, depending on the precise construction of thefacesheets. FIG. 12G illustrates the carry through component 135 securedto one surface of the mixer duct shell 134′. FIG. 12H illustrates thecarry through component 135 with its receptacles 135 b and 135 c securedto a rivet component 135 d sitting in a countersunk opening in a surfaceof the mixer duct shell 134′. FIG. 12I illustrates that when a carrythrough component 135 is used to extend through the honeycomb,preferably it will extend through a point of intersection of a pluralityof the cells of the honeycomb layer 137 a. The intersecting corners of aplurality of cells will be machined out or otherwise removed to form anelongated hole through the honeycomb 137 a that the elongated element135 a of the carry through component 135 may extend.

Referring now to FIGS. 13-17, the structure of the mixer 194 will bedescribed in greater detail. The mixer 194 forms a circumferentialcomponent having a circumferential forward body portion 194 a, attachedin part to an aft portion of the engine turbine case 141′ and also to anaft portion of a core exhaust shroud 141 a′, as shown in FIG. 14. Themixer 194 forward body portion 194 a narrows into a center body tube 194b, within which a second, smaller diameter co-annular tube 194 k islocated. A plurality of radially extending support vanes 194 c extendabout the periphery of the forward body portion 194 a and are attachedto the aft engine turbine frame 166′ (FIG. 8), and form an outer mixerflow path 194 d (FIG. 15); as well as form points of attachment fromwhich the mixer duct shell 134′ can be supported radially offset fromthe mixer 194. A plurality of circumferentially spaced annular cavityvent paths 194 e (FIGS. 13, 14 and 17) are integrally formed in theforward body portion 194 a, circumferentially about the forward bodyportion 194 a, and extend into communication with the center body tube194 b through ports 194 e 1 (FIG. 13c ). This plurality of cavity ventpaths 194 e conducts engine case ventilation flow emanating from a gapbetween the core engine case and the interior surface of the inner fancase shroud in a radially. inward direction flowing into a co-annularflow path 194 b 1 (FIG. 13) between the inner surface of the center bodytube 194 b and the outer surface of the co-annular concentric tube 194k. A plurality of scalloped projecting portions 194 f (FIGS. 13 and 14)are formed circumferentially about the forward body portion 194 a andhelp to form a plurality of inner mixer flow paths 194 g, as bestillustrated in FIGS. 15 and 16. A mixer trough 194 g 1 extendsdownstream of each inner mixer flow path 194 g, as seen particularlywell in FIGS. 13b and 14. The inner mixer flow paths 194 g will bereferred to collectively hereinafter simply as the “inner mixer flowpath 194 g”. A circumferential wall portion 194 h couples to the aftportion of the core engine turbine case 141′ in FIG. 14, while wallportion 194 a 1 is secured to the core exhaust shroud 141 a′. An axiallydisposed vent exit 194 j is formed at an aft end of the center body tube194 b. Within the vent exit 194 j is the concentric inner tube 194 k.

With reference to FIG. 14, the outer mixer flow path 194 d is used forreceiving a pressurized fan exhaust flow 500, which typically has atemperature around 100 degrees Fahrenheit. The annual cavity vent paths194 e (FIG. 17) receive a low pressure flow of engine case ventilationair 502, which typically has a temperature around 600 degreesFahrenheit. A pressurized core exhaust flow 504, typically on the orderof around 1200 degrees Fahrenheit, flows into the inner mixer flow path194 g. Low pressure engine shaft bearing cooling air 506, typically onthe order of around 500 degrees Fahrenheit, flows into an inner area 194i defined by an interior surface of wall portion 194 h.

Referring now to FIG. 18, the directions and mixing of the various flowswill be described in further detail. Ambient air 508 flows around theouter perimeter of the short nacelle 36′. Ambient air 508 also enterswithin an interior area of the short nacelle 36′ and is drawn into theengine 106′ by fan 124′. The ambient air 508 is heated as it flows overan outer perimeter of a core exhaust engine shroud 108′ (FIG. 13B) andbecomes the pressurized fan exhaust flow 500 as it moves toward the aftend of the engine 106′. A first portion 500 a of the pressurized fanexhaust flow 500 is routed through an annular opening formed between anouter surface of the mixer shell 134′ and an inner surface of the aftend of the short nacelle 36′. A second portion 500 b of the pressurizedfan exhaust flow 500 flows into the outer mixer flow path 194 d formedbetween adjacent ones of the radially extending fins 194 c. The lowpressure shaft bearing cooling flow 506 flows out of engine 106′, intothe interior area defined by wall portion 194 h, and is channeledthrough the interior area of the concentric tube 194 k. Engine caseventilation flow 502 is channeled through the integral vent mixer tubes194 e, through ports 194 e 1, and radially inwardly into the annularcavity 194 b 1 formed between the interior surface of the center bodytube 194 b and the outer surface of the concentric tube 194 k. The flows506 and 502 exit separately through the center concentric tube 194 k andthe annular cavity 194 b 1. Both flows have temperatures typicallybetween about 500-600 degrees Fahrenheit, and slightly differentpressures that demand separate exits to prevent one from back pressuringthe other. For some engines the pressures of the two vent flows can besubstantially the same allowing the concentric inner tube 194 k flow tobe mixed within the interior cavity of the center body tube 194 b. Inthat instance, a mixed flow, combined from the two flows 506 and 502,would exit the vent exit 194 j of the center body tube 194 b.

The mixer 194 overcomes deficiencies of previously used mixingapproaches by adding efficient means of conducting the low pressure casevent air through the above defined flow path out the center line of thepressurized higher-velocity mixed exhaust flow. This overcomes theproblem of trapping engine case heat or providing a substantialefficiency loss to, in effect, pump that flow out of the engine 106′through other pneumatic or forced transport means. Mixer 194 providesthe benefit of a simple and efficient transport of the vent cooling air,to keep the engine 106′ cool during all phases of operation where flowis exhausting from the mixer 194.

Another important benefit is the significant thermal cooling on themixer 194 itself that is provided by the integral annular cavity ventpaths 194 e. The vent paths 194 e significantly reduce the temperatureexposure on the leading edge surface 194L (FIG. 17) of the mixer 194that is protruding into the hot pressurized core exhaust flow 504. Thisresults in significantly improved cooling of the mixer device 194, andespecially at the leading edge 194L. This can also significantly reducethe thermal fatigue on the mixer 194 as compared to other conventionalmixer devices. The significantly improved cooling of the mixer 194 mayalso allow a reduction in mixer 194 weight through the use of thinnergauge materials, while still allowing the materials to meet theanticipated thermal stresses that the mixer 194 will experience duringoperation of the jet engine 106′.

With further reference to FIG. 18, the extremely hot pressurized coreexhaust flow 504 flows into the inner mixer flow path 194 g and mixeswith the second flow portion 500 b of the much cooler pressurized fanexhaust flow 500 b, which is flowing radially inwardly between thescalloped projecting portions 194 f toward the center body tube 194 b.This mixing occurs through the turbulent interleaving of two pressurizedconcentric flows (inner hot and outer cool) by a plurality of peaks andtroughs that disperse the hot and cold flows into interleaved(alternating hot and cool) areas confined by the peaks of the mixerlobes 194 f and recessed troughs 194 g 1. This significantly cools theincoming pressurized core exhaust flow 504 to create a flow 512 thatexits the mixer 194 at a temperature typically around 600 to 700 degreesFahrenheit, or close to that of the engine case ventilation flow 502.This has the benefit of significantly reducing the heat “signature” ofthe engine 106′, and may contribute to making the exhaust plume of theengine 106′ less susceptible to heat detection in military aircraftapplications.

Referring now to FIGS. 19 and 20 the difference in how the mixer 194influences thrust reverser operation can be seen. FIG. 19 is anillustration of a prior art thrust reverser being used on a turbofan jetengine with the mixer 194 and the mixer duct shell 134′ of the presentdisclosure. Engine case air does not pump during thrust reverseroperation. With reference to FIG. 20, however, with the mixer 194 andthe mixer duct shell 134′, the engine case ventilation air 502 stillpumps during thrust reverser operation. During this operation an ambientair flow 508 is drawn in through the engine fan exhaust exit,reversedly, and around the gap between the mixer duct shell 134′ leadingedge and the back of the deployed thrust reverser gate. The flow 508then flows into the mixer 194 and passes through the mixer 194, thusmixing with the diffusing core flow leaving the scalloped mixer lobes194 f. The flow diffusion is tempered by the siphoning action of thereverse ambient air flow through the mixer duct resulting in the pumpingof the mixture of the flows and the cooling of temperature and slowingof velocity of the mixture. The much slower velocity, lower temperaturediffused exhaust flow, represented by arrows 512, subtractssubstantially less from the reverse thrust generated by the fan ductflow 500, thus making thrust reverser performance even better than thatfrom the common class of prior art engines operating fan exhaust thrustreversers.

Various implementations of the disclosure can provide comprehensivereduction in both noise and heat radiation over an entire takeoff andapproach operational envelope by virtue of a partial amount of bypassinternally ducted and efficiently mixed with high-velocity hot coreexhaust. This greatly reduces community noise for commercial aircraftand infrared threat susceptibility for military aircraft. Variousimplementations also can provide reduction in plume energy sufficient toeliminate or reduce potential on-board sensor visibility interferencefor countermeasures systems, when looking directions through ownaircraft generated plumes, thereby increasing protection against, e.g.,to man-portable air defense (MANPAD) systems.

An additional benefit is that various implementations also can enablehigh aerodynamic performance and greater cruise range flexibility ataltitudes away from populations and threats, where the feature ofinternal mixing enhances fuel economy. At lower altitudes,implementations in accordance with principles of the disclosure canoffer an ability to maximize thrust performance (allowing increasingfuel capacity or payload lift at takeoff), while keeping noise and heatemissions from the engine plumes well controlled. Modest changes can bemade to outermost short-nacelle surfaces in accordance with principlesof the disclosure while retaining high aerodynamic efficiency, low drag,and excellent propulsive performance.

While various embodiments have been described, those skilled in the artwill recognize modifications or variations which might be made withoutdeparting from the present disclosure. The examples illustrate thevarious embodiments and are not intended to limit the presentdisclosure. Therefore, the description and claims should be interpretedliberally with only such limitation as is necessary in view of thepertinent prior art.

What is claimed is:
 1. An engine assembly, comprising: a core engine; afan driven by the core engine; a short nacelle disposed around the fanand a forward portion of the core engine; a lobed mixer; and a mixerduct shell surrounding the lobed mixer, at least a portion of the mixerduct shell disposed between the short nacelle and an aft portion of thecore engine, the mixer duct shell comprising: a woven compositecontainment liner positioned on an upstream end of the mixer duct shellon an inward-facing surface of the mixer duct shell, the woven compositecontainment liner extending upstream of the lobed mixer; and an acousticliner comprising a honeycomb structure, the acoustic liner positioned ona downstream end of the mixer duct shell on the inward-facing surface ofthe mixer duct shell, the acoustic liner extending downstream of thelobed mixer.
 2. The engine assembly of claim 1, wherein the core engineis covered by a core engine shroud.
 3. The engine assembly of claim 1,wherein the woven composite containment liner is adjacent a leading edgeof the mixer duct shell.
 4. The engine assembly of claim 1, furthercomprising an acoustic lining disposed on an outward-facing surface ofthe mixer duct shell.
 5. An aircraft, comprising: a wing pylon; and anengine assembly coupled to the wing pylon, the engine assemblycomprising: a core engine; a fan driven by the core engine; a shortnacelle disposed around the fan and a forward portion of the coreengine; a lobed mixer; and a mixer duct shell surrounding the lobedmixer, at least of a portion of the mixer duct shell disposed betweenthe short nacelle and an aft portion of the core engine, the mixer ductshell comprising: a woven composite containment liner positioned on anupstream end of the mixer duct shell on an inward-facing surface of themixer duct shell, the woven composite containment liner extendingupstream of the lobed mixer; and an acoustic liner comprising ahoneycomb structure and a facesheet, the acoustic liner positioned on adownstream end of the mixer duct shell on the inward-facing surface ofthe mixer duct shell, the acoustic liner extending downstream of thelobed mixer.
 6. The aircraft of claim 5, wherein the core engine iscovered by a core engine shroud.
 7. The aircraft of claim 5, wherein thehoneycomb structure comprises a ceramic material.
 8. The aircraft ofclaim 7, wherein the honeycomb structure comprises a ceramic matrixcomposite.
 9. The aircraft of claim 8, wherein the facesheet is weldedto the honeycomb structure.
 10. The aircraft of claim 9, wherein thefacesheet comprises titanium.
 11. The aircraft of claim 5, wherein thefacesheet is perforated.
 12. The aircraft of claim 5, further comprisingan acoustic lining disposed on an outward-facing surface of the mixerduct shell.
 13. The aircraft of claim 5, wherein the woven compositecontainment liner is adjacent a leading edge of the mixer duct shell.14. An engine assembly, comprising: a core engine surrounded by a coreengine shroud; a fan driven by the core engine; a short nacelle disposedaround the fan and a forward portion of the core engine shroud, a fanpressurized duct defined between the short nacelle and the forwardportion of the core engine shroud; a lobed mixer; and a mixer ductshell, at least of a portion of the mixer duct shell disposed betweenthe short nacelle and an aft portion of the core engine, the mixer ductshell comprising: a woven composite containment liner positioned on anupstream end of the mixer duct shell on an inward-facing surface of themixer duct shell, the woven composite containment liner extendingupstream of the lobed mixer; and an acoustic liner including a honeycombstructure comprising ceramic, the acoustic liner positioned on adownstream end of the mixer duct shell on the inward-facing surface ofthe mixer duct shell, the acoustic liner extending downstream of thelobed mixer.
 15. The engine assembly of claim 14, wherein the honeycombstructure comprises a ceramic matrix composite.
 16. The engine assemblyof claim 15, further comprising a second acoustic liner disposed on anoutward-facing surface of the mixer duct shell.
 17. The engine assemblyof claim 16, further comprising a third acoustic liner disposed on thecore engine shroud.
 18. The engine assembly of claim 17, wherein thethird acoustic liner comprises a perforated facesheet.
 19. The engineassembly of claim 18, wherein the perforated facesheet comprises a metalalloy.
 20. The engine assembly of claim 18, wherein the perforatedfacesheet comprises a ceramic matrix composite.